CN114237295A - Unconventional flight control technology for high-agility air-to-air missile at large angle of attack - Google Patents
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Abstract
The invention discloses an unconventional flight control technology for an air-to-air missile at a large attack angle, which is different from a method based on control plane, thrust vector and direct force control under the conventional large attack angle. The exciter is arranged near the head of the hollow missile of the slender body, the exciter is utilized to change the relative position of the front body vortex and the surface of the missile, and lateral force and yaw moment are induced, so that the missile is controlled to change the flight direction, and the missile can realize turning at the maximum of 180 degrees. The invention designs a front body vortex control system with sideslip angle detection, which can be independent of a main control system, automatically eliminate the sideslip angle, and can also actively provide yaw moment through the front body vortex when the sideslip angle is needed or the yaw moment is needed to be actively provided during flying, thereby realizing unconventional flight control of the high-sensitivity air-to-air missile large attack angle.
Description
Technical Field
The invention relates to the field of aerospace, in particular to a high-sensitivity unconventional flight control technology for an air-to-air missile at a large angle of attack.
Background
Through the development of many years, the air-to-air missile is launched by a fixed shaft which needs to chase the tail part of an enemy plane until the launching at a small off-axis angle can be realized, and then the realized omnidirectional hitting capability is realized, namely the capability of hitting any angle enemy plane including a rear hemisphere is realized by taking the air-to-air missile as a center. At present, air-to-air missile launching is mainly divided into three stages, namely primary guidance, intermediate guidance and final guidance. In the initial guidance process, the axis and the speed direction of the missile need to be quickly transferred to the vicinity of a target line so as to meet the requirement of guidance by a final guidance method. In order to strike a rear semi-spherical target, the initial guidance of the front-launched missile needs to pass through a rapid turning process, namely an agile turning process of launching the missile over the shoulder. This process can be accomplished by an over-stall maneuver, i.e., a short period of time to rapidly increase the angle of attack, then roll along the speed axis, and then decrease the roll and angle of attack after the end of the turn is reached, as shown in FIG. 1. The missile is pulled up to a large attack angle so as to decelerate and reduce the turning radius. During the whole maneuvering process, the sideslip angle needs to be maintained to be zero, and then an additional yaw moment is needed to eliminate the sideslip angle. However, at high angles of attack, there is a large flow separation, resulting in a significant reduction in the control efficiency of the aerodynamic control surface. Other control methods are needed to achieve flight control at large angles of attack.
The existing control method adopts the compound control of vector thrust and an aerodynamic control surface or the compound control of direct force and the aerodynamic control surface to provide yaw moment to eliminate a sideslip angle. For vector thrust, the structure of the tail nozzle is complicated, the deflection angle of an engine is limited, and the range can be reduced after deflection. And many vector engines at present are two-dimensional vector engines, and cannot provide yaw moment. The direct force control method can increase the quality and the volume of the missile, does not meet the requirement of reducing the volume of the air-to-air missile as far as possible, has high temperature of a direct force nozzle and has higher requirement on the surface material of the missile.
In view of the above, the present invention proposes another control method, using the precursor vortex for control to provide yaw moment. The precursor vortex control is to form a high-level vortex and a low-level vortex by actively controlling the asymmetry of the precursor separation vortex, wherein the low-level vortex generates a low-pressure area, the high-level vortex generates a high-pressure area, a pressure difference is formed, and a lateral force and a yaw moment are generated, and the schematic diagram of the principle is shown in fig. 2 and fig. 3. The principle basis of the invention is high attack angle aerodynamics, and because the flight speed of modern missiles is very high, air-air missiles are often designed into slender rotating bodies for reducing resistance. A pair of separation vortices occurs in the lee of the precursor of an elongate rotating body at a high angle of attack. On the basis, the precursor vortex is actively controlled through the exciter, the precursor vortex is controlled to have an asymmetric configuration, a lateral force is induced, and the lateral force generated by the precursor can generate enough control moment because the missile precursor is far away from the center of mass of the missile. At the same blowing rate, the precursor vortex control can generate control force and control torque which are three orders of magnitude greater than the thrust of the jet blowing itself. The method is a very efficient method for controlling the air-to-air missile under the large attack angle.
Disclosure of Invention
1: the precursor vortex control system consists of:
the invention relates to a system for automatically judging a sideslip angle when a missile finishes agile turning in large-attack-angle flight, then controlling an exciter according to the sideslip angle, constructing an asymmetric precursor vortex, inducing a yaw control moment and automatically eliminating the sideslip angle. The precursor vortex control system may also be made active to generate a control yaw moment when the flight requires a sideslip angle.
The overall process flow is shown in fig. 4, wherein the upper branch line comprises a main flight control system of the missile by a main controller and a main control effector, and the lower branch line comprises a front vortex control system by a sideslip angle detection module, a control gain module and a front vortex control exciter module. The diamond shaped blocks in fig. 4 are a decision block that provides the synthetic jet device control signal from the control gain device of the down leg when the sideslip angle is not required for flight, and the control signal from the main controller to the precursor vortex control actuator when the sideslip angle is required for flight or when the yaw moment is actively provided.
When a maneuver is performed that does not require a sideslip angle, the front body vortex control system and the aircraft main control system are decoupled from each other without increasing the complexity of the main control system. In the primary flight control system, the sideslip angle can be defaulted to zero, which can provide a simplification to the primary flight control system. The main control system provides the missile with the control torque required for normal control except sideslip elimination, and the front vortex control system is responsible for eliminating sideslip. In the flight process, a sideslip angle detection module detects a sideslip angle, a sideslip angle signal is transmitted to a control gain module to be converted into a control signal, and the control signal is transmitted to a front body vortex control exciter, so that the front body vortex control exciter generates a yawing moment in the opposite direction. The method is specifically divided into three cases:
when no sideslip exists, beta is equal to 0, the exciter does not work, the left front body vortex and the right front body vortex are symmetrical, no lateral force is generated, and no yaw moment is generated. The sideslip angle remains unchanged.
When the right side slides, beta is more than 0, the exciter generates asymmetric vortex, the left side is a low-position vortex, and the right side is a high-position vortex, so that the lateral force in the positive direction of the y axis is generated, and the yaw moment in the positive direction of the z axis is generated. Eliminating right sideslip.
When the left side slips, beta is less than 0, so that the exciter generates an asymmetric vortex, the left side is a high-level vortex, the right side is a low-level vortex, the y-axis negative-direction lateral force is generated, and the z-axis negative-direction yawing moment is generated. Eliminating left sideslip.
When a maneuver requiring a sideslip angle is performed or a yaw moment is actively provided, a control signal is directly provided to the control actuator of the precursor vortex by the main controller, and the controller for separating the precursor vortex is controlled in combination with other control effectors in the same manner as the conventional flight control, which will be shown in the example section and not described herein.
The control method shown in fig. 4 is the unconventional flight control method of the high-agility air-to-air missile large angle of attack.
2: application example:
2.1: application example overview:
in order to introduce the method more clearly and prove that the method is really usable, a dynamic simulation model in the form of a graph 4 of a geometric model in a graph 7 is established as an example, the established dynamic simulation model is subjected to herbst maneuver simulation, and whether the herbst maneuver simulation can be completed under the control method shown in the graph 4 to complete agile turning is checked.
All the contents in the second section belong to the application example part, the algorithm formula is not the content of the invention, but the invention is only an example, and the method is proved to be effective. And is replaceable in practical application.
2.2: establishment of precursor vortex control actuator:
the precursor vortex control exciter module is formed by arranging a pair of blowing holes which are symmetrical along a longitudinal symmetry plane at the leeward position of the missile precursor, wherein the axial position of the blowing holes is the starting position of the precursor vortex, and the circumferential position of the blowing holes is near the leeward longitudinal symmetry plane (between +/-45 and +/-30), as shown in figure 6. The synthetic jet flow is used for controlling the front body vortex of the air-air missile, the distance between the front body vortex on one side and the surface of the missile is changed by changing the blowing momentum of the blowing hole during control, so that the front body vortex on the blowing side is changed into a high-position vortex from a low-position vortex, a pressure difference is formed between the front body vortex and the low-position vortex on the non-blowing side, a lateral force and a yawing moment are generated, and the aim of controlling the yawing moment to finish the agile turning of the missile is fulfilled.
Defining the blowing momentum of the blowing hole as
Wherein u isjIs the outlet velocity, s, of the micro-blowing holejIs the area of the micro-blowing hole u∞For incoming wind speed, D is the elongated body diameter. And defining the blowing momentum of the right hole as positive and the blowing momentum of the left hole as negative. The right side hole is a hole in the negative direction of the y axis below the body axis. The total blowing momentum is as follows
Cμ＝Cμ-right-Cμ-left (2)
The current sideslip angle of the missile will determine the blow-off magnitude, and the magnitude of the blow-off momentum will affect the location of the precursor vortex, for three cases, where K is the gain.
When there is no sideslip, beta is 0, CμAnd when K beta is 0, no air blowing hole blows air, the left and right precursor vortices are symmetrical, no lateral force is generated, and no yawing moment is generated. The sideslip angle remains unchanged.
Beta is more than 0 when sliding on the right side, CμThe left side is a low vortex, and the right side is a high vortex, so that the lateral force in the positive direction of the y axis is generated, and the yawing moment in the positive direction of the z axis is generated. Eliminating right sideslip.
Beta < 0, C when left side slippingμAnd when K beta is less than 0, the left side is blown by the blowing hole, the left side is a high-level vortex, and the right side is a low-level vortex, so that the y-axis negative-direction lateral force is generated, and the z-axis negative-direction yawing moment is generated. Eliminating left sideslip.
2.3: establishing an application example missile dynamics model:
the missile module in fig. 4 consists of an aerodynamic model and a dynamic model. The stress and moment conditions of the missile under the conditions of different flight states and control quantity input are obtained by calculating the aerodynamic force model and CFD numerical values. And then modeling the obtained pneumatic data, wherein the modeling method uses a radial basis function neural network for modeling. The aerodynamic influence factor is shown in equation 3:
Ci＝Ci(α,β,δy+,δy-,δz+,δz-,Cu,P,Q,R) i＝X,Y,Z,L,M,N (3)
the independent variables are, in order, an attack angle, a sideslip angle, a y-axis positive direction elevator deflection angle, a y-axis negative direction elevator deflection angle, a z-axis positive direction rudder deflection angle, a z-axis negative direction rudder deflection angle, an air blowing momentum (a control method to be researched in the text), a rolling angular velocity, a pitch angular velocity and a yaw angular velocity. The dependent variable is axial force coefficient, lateral force coefficient and normal force coefficient in sequence. And then divided into static and dynamic parts as in equation 5. The static part is directly measured by CFD numerical calculation and then modeled by a radial basis function neural network, and the dynamic part is decomposed by a Kalvist method as shown in a formula 6, wherein Pmod QmodRmodAngular velocities, ω, along the x, y, z axes of the body axis after decomposition, respectivelyssIs the angular velocity along the velocity vector after the decomposition. And then, dynamic numerical calculation is carried out by utilizing CFD (computational fluid dynamics), and numerical calculation of vibration along three axes and rolling along a speed axis is respectively carried out, so that the numerical value of each part on the right side of the equal sign in the formula 6 is respectively obtained. And obtaining a pneumatic database, and finally establishing and simulating the equation by using a neural network. Finally, the force coefficient and the moment coefficient of the missile in different flight states are obtained.
Ci＝Cistatic(α,β,δy+,δy-,δz+,δz-,Cu)+Cidynamic(α,β,P,Q,R) (5)
The kinetic model is built from the system of equations of full-scale motion for the missile, as shown in equation 8. Wherein X, Y and Z are components of aerodynamic force in an airplane system, and a coordinate system is selected as an American coordinate system. L, M, N are the components of the aircraft aerodynamic moment on the airframe. Calculated from equation 7.
1/2 rho V2Is dynamic pressure, S is the wing area of the airplane, b is the wingspan,
is the mean aerodynamic chord length.
Wherein phi, theta and psi are respectively the yaw angle, pitch angle and roll angle of the aircraft; u, v and w are speeds along x, y and z axes under the body axis respectively; p, q and r are angular velocities along x, y and z axes under the body axis respectively; x is the number ofv,yv,zvIs the position under the ground coordinate system; m and g are respectively mass 220kg and gravity acceleration 9.8N/s2；Ix,Iy,IzRespective moments of inertia along three axes, respectively 0.788m4，163.607m4，163.607m4；IxzThe product of inertia is zero because the missile is axisymmetric. The entire system of equations satisfies the assumption of a non-spinning, flat earth, aircraft with a plane of symmetry.
And integrating the formula to obtain the flight states of the missile under different aerodynamic force aerodynamic moments. Thus, the establishment of a missile simulation model is completed, and the flight state of the missile is completely known under the condition of determining the initial value and the control quantity.
2.4: establishment of application instance master control system
The main control effector in fig. 4 consists of the four tail control surfaces of the missile and the vector motors providing the pitch and roll moments. The main controller consists of a dynamic inverse control system and a base sequence control allocation algorithm.
The dynamic inverse control algorithm obtains the expected control torque through the expected attack angle, sideslip angle and speed roll angle, and needs to be divided into two steps. First, a desired angular velocity is obtained by the attitude input, and a desired control torque is obtained from the desired angular velocity.
According to the differential equation:
wherein [ Fx,Fy,Fz]TThe stress of each shaft under the wind shaft system is shown, and V is the sum speed.
Wherein, the derivatives of the attack angle, the sideslip angle and the speed and roll angle can be determined by the following modes:
wherein [ k ] isα,kβ,kμ]TNeeds to be determined according to actual conditions, and is taken as [5,0,5 ] in the text]TWherein since the desired sideslip angle defaults to zero and the sideslip angle of the true flight condition is eliminated by the precursor vortex control system, the sideslip angle of the true flight condition is also zero by default here, and naturally the difference between two zeros is also zero, so kβ＝0。
K when sideslip angle is required or active provision of yaw moment is requiredβAnd not equal to 0, determined according to actual conditions.
The desired angular velocity can be found by transforming the differential equation:
the desired control torque is obtained by conversion according to the rotational dynamics differential equation:
wherein the desired angular velocity derivative is:
wherein [ k ] isp,kq,kr]T＝[2,2,2]. The specific coefficient needs to be determined according to actual conditions, the expected moment can be obtained at the moment, the moment obtained is the expected resultant moment borne by the aircraft and is not the required control moment, so the moment generated by the current missile in an uncontrolled way needs to be reduced, and the uncontrolled moment [ L ]b,Mb,Nb]TIncluding the moment created by the projectile, the missile wing, airflow disturbances, etc. The required control torque is therefore:
up to this point, the required control torque has been determined according to the desired attitude.
In the example, the main control effector is controlled in two modes, namely an aerodynamic control surface mode and a vector engine mode. However, in flight, because the vector engine has large deflection fuel consumption, the sequence of control is that the pneumatic control surface and the front vortex are firstly used, and if the two control modes can not achieve the required control torque, the vector engine is started. The following equation:
yc＝yc1+yc2 (15)
wherein y iscRequired control moment [ L ] determined for the previous sectionec,Mec,Nec]T，yc1Control moment, y, generated for the fore-body vortex and the aerodynamic control surfacec2Is the control torque generated by the vector motor. In general, yc2Is zero.
Knowing the required control torque, it is also necessary to find the opening of each control effector to produce the required control torque, i.e. to solve the following equation:
wherein y isc1＝[Lec1,Mec1,Nec1]TB is the control efficiency matrix and u is the opening vector of the control effector, where u is [ δ ═ in this contexty+,δy-,δz+,δz-]TThe right end of the equation is the tail deflection angle in the positive y-axis direction, the tail deflection angle in the negative y-axis direction, the tail deflection angle in the positive z-axis direction, and the tail deflection angle in the negative z-axis direction, respectively.
When the yaw moment needs to be actively provided, the blowing momentum u ═ delta needs to be addedy+,δy-,δz+,δz-,Cu]T。umin,umaxThe upper and lower limits of each control effector, in this example the upper and lower limits of the four tail angles of declination, are plus or minus 25 degrees.
The missile is composed of four control surfaces on the tail, four variables are total, and five control surfaces are needed to actively provide yaw moment. The control effector is redundant, i.e. the dimension of u is greater than y in the formulac1The equation may be multi-solution, and may be solution-free after the upper and lower bounds of the control effector are added. Additional conditions and control allocation algorithms are therefore required to determine the degree of opening of the individual control effectors given the required control torque. The method of base sequence allocation is adopted here, and the method is complicated and is not the key point of the present invention, and reference may be made to the documents Calibration and compensation of near-field scan measurements, which are not described herein again.
The selection and establishment of the master control effector and the master controller is completed.
2.5: application instance simulation results
At this point, the establishment of all modules of the integral missile simulation model shown in fig. 4 is completed.
Under the model, the flight state of the missile can be completely determined by determining the initial value and the target flight state. The initial values are selected as follows:
the target flight state is an over-stall maneuver, mainly comprises three parts, namely a pulled attack angle and a rolling along a speed axis, when the azimuth angle of the flight path is close to 180 degrees, the attack angle and the rolling angle of the speed are reduced, the flight path returns to a level flight state, and turning around is completed. As shown in fig. 8 for control inputs, and fig. 9 for simulation results.
As can be seen from the simulation results, the sideslip angle is controlled within 1.5 degrees, and the over-stall maneuver is completed. In the case that the sideslip angle is defaulted to zero in the main control system, the front vortex control system is independent of the main control system, and the missile completes the agile turn of the first stage of over-the-shoulder launching, the front vortex control system can provide enough and accurate yaw moment to eliminate the sideslip angle.
3: compared with the prior art, the method has the following advantages:
(1) the structure and mass of the front body vortex control system is much smaller than the direct force and vector engine for the same yaw moment.
(2) In a motor without a sideslip angle, the front vortex control system can automatically eliminate sideslip, is independent of the main control system of the missile, and plays a role in independent work, so that the complexity of the main control algorithm cannot be increased.
(3) In maneuvers requiring a sideslip angle, the control system of the precursor vortex may also actively provide a yaw moment, which adds to the other control effector effects, resulting in an increased control moment.
Drawings
FIG. 1herbst schematic diagram of the maneuver
FIG. 2 active control generating asymmetric precursor vortices schematic
FIG. 3 three-dimensional schematic diagram of actively controlled generation of asymmetric precursor vortices
FIG. 4 is a flow chart of a high-agility air-to-air missile large-attack-angle unconventional flight control method
FIG. 5 flow chart of precursor vortex control system
FIG. 6 is a schematic view of the position of the air blowing hole
FIG. 7 geometric model diagram of a simulation model
FIG. 8 flight simulation example control inputs
FIG. 9 flight simulation example results
Detailed Description
The invention can be implemented according to the following implementation method:
1. firstly, wind tunnel experiment or numerical calculation is carried out on the condition of large attack angle of the agile missile, the position of a precursor vortex is determined, and an exciter is arranged at the initial position of the precursor vortex.
2. And determining the lateral force and yaw moment which can be generated by different exciter openness under different attack angles through wind tunnel tests or numerical calculation. The range in which the actuator opening degree is linearly related to the yaw moment, and the linear correlation coefficient are determined.
3. The flight control system as shown in fig. 4 was redesigned. The system comprises a main flight control system of the missile and a precursor vortex flight control system. The main flight control system consists of a main controller and a main control effector. The precursor vortex control system consists of a sideslip angle detection module, a control gain module and a precursor vortex control actuator, as shown in fig. 5. The precursor vortex control system is connected in parallel with the main control system of the missile, where the sideslip angle is defaulted to zero when not needed.
4. And performing flight simulation on the missile, and determining the gain of the control gain module to enable the precursor vortex to generate a yaw moment meeting the requirement to complete maneuvering.
Various modifications and alterations may be made by those skilled in the art without departing from the spirit and scope of the invention. For example, in the control method shown in fig. 4, the main control system can be freely selected according to different missiles, and is not limited to the control method of dynamic inverse and base sequence in the example, and the main control effector is not limited to an aerodynamic control surface and a vector engine; in precursor vortex control systems, the choice of actuator is not limited to synthetic jets. Such modifications and variations are within the scope of the invention as defined by the appended claims.
Claims (5)
1. The unconventional flight control technology for high-agility air-to-air missile at large attack angle is characterized in that an exciting device is arranged near the head of a missile forebody and serves as a forebody vortex separation controller.
2. An unconventional flight control technology for high-agility air-to-air missile at a large angle of attack is characterized in that whether a sideslip angle is needed or not is judged according to a target flight state.
3. The unconventional flight control technology for the high-agility air-to-air missile at the large attack angle as claimed in claim 2 is characterized in that when the missile does not need the sideslip angle, the missile detects the sideslip angle of the missile through a sideslip angle detection system under the high attack angle, so that a controller of a precursor separation vortex generates yawing moments in opposite directions to achieve the aim of eliminating sideslip.
4. The unconventional flight control technology for the high-agility air-to-air missile at the large attack angle as claimed in claim 2, is characterized in that when the missile needs a sideslip angle, a controller control signal provided by the main control system to the precursor separation vortex actively provides a yaw moment.
5. The unconventional flight control technique for the high-agility air-to-air missile at the large attack angle as claimed in claim 3, wherein the yaw moment is provided by the front vortex control system independently from the main control system and the sideslip angle in the main control system is set to zero by default.
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| CN116560412A (en) * | 2023-07-10 | 2023-08-08 | 四川腾盾科技有限公司 | Test flight planning method for verifying maximum flat flight speed index of low-speed unmanned aerial vehicle |
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